The invention relates to rotorcraft rotors comprising a hub driven in rotation about an axis of rotation by a drive shaft or outlet shaft from a power transmission gearbox, and also having at least three blades fixed to the hub via appropriate hinges, in particular via respective laminated spherical abutments dedicated to each of the blades, with inter-blade dampers interconnecting each pair of two adjacent blades.
On the assumption that each blade is inserted in a hub in fixed manner, the resulting rotor would be a rigid rotor. When hovering, the distribution of aerodynamic forces along a blade leads to a bending moment distribution presenting a value that is very large at the root of the blade. While flying in translation, the so-called “retreating” blade carries a greater load than the so-called “receding” blade because of the different air speeds of the blades, as described in greater detail below.
Consequently, the resultant of the aerodynamic forces exerted on a blade does not have the same value at each azimuth position, nor does it have the same point of application: the fixed-end moment at the blade root is thus high and varying, leading to alternating stresses that give rise to a fatigue phenomenon that is harmful to materials. In addition, the resultant of the aerodynamic forces on all of the blades is no longer directed along the axis of the rotor, thereby creating a roll moment that increases with increasing speed and that can make it difficult to balance forces when flying in translation.
In order to remedy those drawbacks, it is known to hinge blades on the hub about an axis, that is perpendicular to the drive shaft and that is designated “vertical flapping axis”, corresponding to an hinge for vertical flapping that can take up arbitrary orientation forces but that can under no circumstances take up a moment. Consequently, if the blade is hinged to the hub, the moment at its point of attachment is zero. In order to enable a blade to be in equilibrium, the centrifugal forces holding up the blade after it has risen a certain amount, causing conicity a0 to appear.
Under such conditions, there is no longer any large roll moment when flying in translation, and furthermore the blades no longer rotate in a plane, but their outer tips describe a widely-open cone. In practice, the flapping axis is then located not on the axis of rotation but is offset therefrom by a distance a, known as its eccentricity.
In order to support a helicopter in its various configurations, it should also be recalled that it is necessary to be able to control the lift provided by the rotor and to cause it to vary. That is why a pitch hinge is provided, of axis that is substantially parallel to the span of the corresponding blade. This further degree of freedom serves to control the lift of the blade by acting on a general pitch control, and also to cause pitch to vary cyclically, thus enabling the plane of rotation of the blades to be controlled so that they then describe a cone having a virtual axis that no longer coincides with the drive axis: the resultant of the forces applied to the hub changes in direction together with the plane of the rotor. This leads to moments being generated about the center of gravity of the helicopter, thus enabling it to be piloted.
As mentioned above, the plane of rotation of the blades may be different from the plane perpendicular to the drive shaft. Under such conditions, it is necessary in particular for each blade to be hinged in drag since the end of each blade is at a distance from the rotor shaft that varies, as explained more precisely below with reference to Coriolis forces. Otherwise, inertial forces would necessarily appear, generating reciprocating bending moments in each blade in its own plane. Such a drag hinge is provided by hinging a blade about a drag axis that is substantially parallel to the rotor axis, and consequently substantially perpendicular to drag forces. To enable such a blade to be driven by the drive shaft, it is naturally essential for the drag hinge to be far enough away from the rotor axis for the moment due to centrifugal forces to balance the moment due to drag and inertial forces, which means that the drag axis must be offset for eccentric by an amount e, and for this to be achieved without the so-called “drag” angle δ being too great.
Consequently, the blades of a hinged rotor of a rotary wing aircraft, in particular of a helicopter, can perform the following four kinds of rotary motion:
i) rotation about the rotor axis;
ii) pivoting about a vertical flapping, axis made possible by the hinge for vertical flapping;
iii) pivoting about the drag axis, also referred to as the horizontal flapping axis, made possible by the horizontal movement hinge also known as the drag hinge; and
iv) pivoting about the blade axis made possible by a pitch hinge (not specific to hinged rotors).
As described in patent FR 2 497 173, for example, three of the above kinds of rotary motion II, III, and IV can be made possible by using a single member such as a laminated spherical abutment, like that used in SUPER PUMA MKII or NH90 helicopters made by the Applicant, in which such a member allows the blade to move in flapping, in drag, and in pitch.
Nevertheless, the oscillations of each blade about its drag axis can become coupled in unstable manner with the movements or the elastic deformation modes of the fuselage, in particular the oscillations of a helicopter standing on the ground on its landing gear: this is at the origin of the phenomenon known as “ground resonance” that can be dangerous for the aircraft when the natural frequency of the oscillations of the blades about their drag axis is close to one of the natural frequencies of the oscillations of the aircraft on its landing gear.
Remedies for that phenomenon consist in introducing damping on the drag axes, in particular by means of a viscous or dry damper device, or indeed by introducing stiffness with the help of blade spacing cables optionally associated with dampers, as in the ALOUETTE helicopter made by the Applicant.
A function analogous to that of the blade spacing cables is provided by resilient inter-blade connections. In practice, that is done by placing a damper between each pair of two adjacent blades, with the fastenings of such a damper to each of two adjacent blades being at equal distances from the rotor center, i.e. at identical radii relative to the rotor center.
Such dampers include resilient return means of determined stiffness and damping for combating resonant phenomena, in particular ground resonance and also resonance in the drive system that can also appear, in particular on board helicopters.
When rotor blades are excited in drag, the blades move away from their equilibrium position and can be distributed unequally in the circumferential direction, thereby creating an unbalance due to the center of gravity of the rotor being displaced away from its axis of rotation. In addition, blades that are moved away from their equilibrium position oscillate about said position with a frequency ωδ, which is the natural frequency of the blades in drag, also referred to as the first drag mode, or the resonant drag mode.
If Ω is the frequency of rotation of the rotor, it is known that the fuselage of the helicopter is thus excited at the frequencies |Ω±ωδ|.
When standing on the ground via its undercarriage, the fuselage of a helicopter constitutes a system comprising a mass suspended above the ground by a spring and a damper in each undercarriage. The fuselage resting on its landing gear thus has its own natural modes of vibration in roll and in pitch. There is a risk of instability on the ground when the frequency at which the fuselage on its undercarriage is excited is close to the natural oscillation frequency |Ω+ωδ| or |Ω−ωδ|, which corresponds to the phenomenon known as ground resonance. To avoid instability, it is known to seek firstly to avoid passing through these frequencies, and if that cannot be avoided, it is necessary to damp the fuselage on its undercarriage sufficiently and also to damp the blades of the main rotor in terms of their drag motion.
Consequently, the stiffness of the drag dampers of the blades of a main rotor needs to be selected so that the natural frequency of the blades in drag lies outside a potential ground resonance zone, while simultaneously having sufficient damping since, while the speed of rotation of the rotor is passing through a critical speed, both as its speed rises and also as its speed falls, the blades need to be damped sufficiently to avoid entering into resonance.
That is why drag dampers with resilient return means of determined stiffness are also known as frequency adapters.
In general, the stiffness of the damper introduces equivalent angular stiffness opposing the angular movements of the blade relative to the hub about its drag axis. The frequency of the resonant mode of the blades in drag can thus be increased so as to keep that frequency away from the two above-mentioned resonance phenomena.
The equivalent angular stiffness is proportional to the square of the lever arm between the damper and the drag axis of the blade, i.e. the distance between the drag axis and the axis passing through the centers of two ball joints of the damper.
Compared with a conventional configuration in which the dampers are interposed between each blade and the rotor hub, configuring dampers in an inter-blade position serves to increase the lever arm between the dampers and the drag axes of the blades, and also serves to cause two dampers to act on each blade in order to avoid ground resonance. The stiffness of each damper can be limited accordingly, and a resulting advantage is a lower level of static force introduced by fitting each damper as an inter-blade adapter. This configuration is thus very favorable for combating ground resonance.
The invention relates in particular to improving rotorcraft rotors as described in patents FR 2 630 703 and U.S. Pat. No. 4,915,585 relating to a rotor head comprising firstly blades connected to the drive hub by hinges suitable for vertical flapping, for drag, and for pitch, in particular laminated spherical abutments as mentioned above, and also resilient return inter-blade ties with incorporated damping, in compliance with the above description.
Those documents describe a rotor in which each blade is fastened to the hub by a sleeve whose ends form yokes each having two mutually-facing and spaced-apart lugs. Each inter-blade damper is fastened to two adjacent blades via two respective ball joints. Each sleeve is fastened to the rotor hub via a laminated spherical abutment and receives two fastener ball joints for two dampers, respectively. These ball joints are centered on the pitch axis of the blade or they are situated in the immediate vicinity of said pitch axis, and they are also fastened between the two branches of the sleeve, outside the center of the pitch, flapping, and drag hinge compared with the hub, corresponding to a laminated spherical abutment.
According to those patents, it is proposed to offset the two ball joints fastened to a sleeve radially along the pitch axis, and where appropriate to offset them laterally on either side of said axis, with these offsets being as small as possible. The two ball joints are centered substantially in the plane of rotation of the rotor.
Such a disposition for the point where the drag dampers are fastened to the sleeve holding the blades to the hub is for the purpose of completely decoupling phase-shifted angular drag movements from pitch and flapping angular movements.
The present invention has the same object.
In practice, it has been found that the disposition of the two ball joints fastening the drag damper as close as possible to each other and to the pitch-changing axis of the blade leads to structures that are complex and that produce results that are disappointing in terms of decoupling pitch, drag, and flapping oscillations.